1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a large air cooled blade in an industrial gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine with multiple rows or stages of stator vanes and rotor blades that are exposed to the hot gas flow passing through the turbine. Because of the extreme hot temperature of the gas flow, the turbine airfoils (both vanes and blades) require cooling to prevent thermal damage and to allow for higher turbine inlet temperatures that result in higher engine efficiencies. The turbine inlet temperature is limited to the first stage airfoils ability to withstand the high temperatures. The maximum useful temperature that an airfoil can be exposed to and operate according to design is based upon the material properties and the amount of cooling that can be produced for the airfoil.
A large industrial gas turbine (IGT) engine is used to produce electrical power and typically has four stages of stator vanes and rotor blades in the turbine section. The first stage rotor blade has an airfoil section that is less than one foot in spanwise length. The last stage rotor blade can be three feet in airfoil length or longer. These large IGT rotor blades can require cooling when the gas flow temperature at the latter stages is high enough to cause thermal degradation of the airfoils. Because of the larger lengths of these airfoils, the blade will require a large amount of twist from the platform to the blade tip. Prior art large turbine rotor blades are cooled by drilling radial holes into the blade from the blade tip and root sections. Limitations of drilling a long radial hole from both ends of the airfoil increases for a large and highly twisted blade airfoil. A reduction of the available airfoil cross sectional area for drilling radial holes is a function of the blade twist. Higher airfoil twist yields a lower available cross sectional area for drilling radial cooling holes. Cooling of the large and highly twisted blade by this manufacturing process will not achieve the optimum blade cooling effectiveness. U.S. Pat. No. 6,910,864 issued to Tomberg on Jun. 28, 2005 and entitled TURBINE BUCKET AIRFOIL COOLING HOLE LOCATION, STYLE AND CONFIGURATION shows this prior art blade with radial cooling holes.